RAE 102 AIRFOIL (rae102-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: RAE 102 AIRFOIL (rae102-il) Reynolds number: 100,000 Max Cl/Cd: 36.66 at α=4° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rae102-il-100000-n5.txt Download as CSV file: xf-rae102-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: RAE 102 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.500 -0.7018 0.09215 0.08700 -0.0127 1.0000 0.0248
-11.250 -0.7174 0.08414 0.07897 -0.0182 1.0000 0.0246
-11.000 -0.7362 0.07681 0.07158 -0.0234 1.0000 0.0244
-10.750 -0.7545 0.07085 0.06552 -0.0272 1.0000 0.0242
-10.500 -0.7751 0.06563 0.06017 -0.0297 1.0000 0.0242
-10.250 -0.7938 0.06147 0.05585 -0.0302 1.0000 0.0241
-10.000 -0.8111 0.05795 0.05215 -0.0289 1.0000 0.0240
-9.750 -0.8255 0.05445 0.04839 -0.0268 1.0000 0.0240
-9.500 -0.8346 0.05068 0.04429 -0.0248 1.0000 0.0241
-9.250 -0.8375 0.04714 0.04037 -0.0227 1.0000 0.0243
-9.000 -0.8361 0.04369 0.03649 -0.0207 1.0000 0.0245
-8.750 -0.8302 0.04057 0.03288 -0.0187 1.0000 0.0248
-8.500 -0.8207 0.03733 0.02931 -0.0170 1.0000 0.0254
-8.250 -0.8068 0.03505 0.02685 -0.0158 1.0000 0.0264
-8.000 -0.7906 0.03343 0.02505 -0.0146 1.0000 0.0280
-7.750 -0.7731 0.03172 0.02309 -0.0134 1.0000 0.0300
-7.500 -0.7537 0.02969 0.02071 -0.0121 1.0000 0.0317
-7.250 -0.7328 0.02784 0.01857 -0.0109 1.0000 0.0330
-7.000 -0.7124 0.02599 0.01659 -0.0097 1.0000 0.0347
-6.750 -0.6932 0.02478 0.01538 -0.0087 1.0000 0.0379
-6.500 -0.6729 0.02366 0.01415 -0.0074 1.0000 0.0417
-6.250 -0.6527 0.02252 0.01284 -0.0060 1.0000 0.0445
-6.000 -0.6357 0.02129 0.01165 -0.0044 1.0000 0.0488
-5.750 -0.6165 0.02046 0.01074 -0.0030 1.0000 0.0554
-5.500 -0.5994 0.01946 0.00975 -0.0013 1.0000 0.0619
-5.250 -0.5813 0.01863 0.00891 0.0004 1.0000 0.0724
-5.000 -0.5635 0.01778 0.00813 0.0020 1.0000 0.0885
-4.750 -0.5461 0.01692 0.00738 0.0036 1.0000 0.1158
-4.500 -0.5298 0.01593 0.00671 0.0053 1.0000 0.1716
-4.250 -0.5160 0.01474 0.00612 0.0072 1.0000 0.2790
-4.000 -0.5038 0.01373 0.00587 0.0098 1.0000 0.4232
-3.750 -0.4870 0.01335 0.00578 0.0120 1.0000 0.5117
-3.500 -0.4688 0.01315 0.00566 0.0140 1.0000 0.5649
-3.250 -0.4504 0.01301 0.00559 0.0160 1.0000 0.6053
-3.000 -0.4321 0.01292 0.00555 0.0179 1.0000 0.6393
-2.750 -0.4008 0.01287 0.00554 0.0173 0.9926 0.6738
-2.500 -0.3644 0.01286 0.00551 0.0157 0.9825 0.7055
-2.250 -0.3286 0.01283 0.00549 0.0143 0.9719 0.7313
-2.000 -0.2928 0.01278 0.00540 0.0128 0.9610 0.7528
-1.750 -0.2552 0.01273 0.00531 0.0110 0.9511 0.7696
-1.500 -0.2165 0.01268 0.00522 0.0089 0.9415 0.7848
-1.250 -0.1808 0.01262 0.00514 0.0074 0.9294 0.7993
-1.000 -0.1447 0.01256 0.00505 0.0059 0.9172 0.8132
-0.750 -0.1084 0.01252 0.00499 0.0043 0.9049 0.8269
-0.500 -0.0722 0.01248 0.00494 0.0028 0.8926 0.8406
-0.250 -0.0362 0.01246 0.00490 0.0014 0.8799 0.8542
0.000 0.0000 0.01245 0.00489 0.0000 0.8672 0.8672
0.250 0.0362 0.01246 0.00490 -0.0014 0.8542 0.8800
0.500 0.0722 0.01248 0.00494 -0.0028 0.8406 0.8926
0.750 0.1084 0.01252 0.00499 -0.0043 0.8269 0.9050
1.000 0.1446 0.01256 0.00505 -0.0059 0.8132 0.9172
1.250 0.1807 0.01262 0.00513 -0.0074 0.7992 0.9294
1.500 0.2165 0.01268 0.00522 -0.0089 0.7848 0.9415
1.750 0.2551 0.01273 0.00531 -0.0110 0.7696 0.9512
2.000 0.2928 0.01278 0.00540 -0.0128 0.7528 0.9611
2.250 0.3286 0.01283 0.00548 -0.0143 0.7313 0.9720
2.500 0.3643 0.01286 0.00551 -0.0157 0.7057 0.9827
2.750 0.4008 0.01287 0.00554 -0.0173 0.6737 0.9928
3.000 0.4317 0.01292 0.00555 -0.0179 0.6393 1.0000
3.250 0.4500 0.01301 0.00559 -0.0159 0.6056 1.0000
3.500 0.4684 0.01315 0.00566 -0.0139 0.5654 1.0000
3.750 0.4865 0.01335 0.00577 -0.0119 0.5109 1.0000
4.000 0.5034 0.01373 0.00586 -0.0097 0.4235 1.0000
4.250 0.5157 0.01474 0.00611 -0.0072 0.2792 1.0000
4.500 0.5295 0.01592 0.00671 -0.0053 0.1723 1.0000
4.750 0.5458 0.01691 0.00737 -0.0036 0.1161 1.0000
5.000 0.5632 0.01778 0.00812 -0.0020 0.0885 1.0000
5.250 0.5811 0.01862 0.00890 -0.0003 0.0726 1.0000
5.500 0.5992 0.01946 0.00975 0.0013 0.0620 1.0000
5.750 0.6163 0.02046 0.01073 0.0030 0.0554 1.0000
6.000 0.6356 0.02128 0.01165 0.0044 0.0488 1.0000
6.250 0.6527 0.02252 0.01284 0.0060 0.0445 1.0000
6.500 0.6729 0.02366 0.01414 0.0074 0.0416 1.0000
6.750 0.6932 0.02477 0.01537 0.0086 0.0378 1.0000
7.000 0.7125 0.02600 0.01659 0.0097 0.0347 1.0000
7.250 0.7330 0.02784 0.01857 0.0108 0.0330 1.0000
7.500 0.7539 0.02971 0.02073 0.0120 0.0316 1.0000
7.750 0.7733 0.03172 0.02309 0.0133 0.0300 1.0000
8.000 0.7910 0.03342 0.02504 0.0145 0.0280 1.0000
8.250 0.8072 0.03501 0.02680 0.0157 0.0263 1.0000
8.500 0.8211 0.03735 0.02933 0.0169 0.0254 1.0000
8.750 0.8307 0.04060 0.03290 0.0185 0.0248 1.0000
9.000 0.8368 0.04372 0.03650 0.0205 0.0245 1.0000
9.250 0.8385 0.04712 0.04035 0.0226 0.0243 1.0000
9.500 0.8346 0.05080 0.04443 0.0247 0.0241 1.0000
9.750 0.8266 0.05445 0.04840 0.0266 0.0241 1.0000
10.000 0.8130 0.05794 0.05213 0.0286 0.0241 1.0000
10.250 0.7958 0.06144 0.05580 0.0300 0.0241 1.0000
10.500 0.7768 0.06562 0.06015 0.0296 0.0242 1.0000
10.750 0.7571 0.07070 0.06537 0.0272 0.0243 1.0000
11.000 0.7362 0.07708 0.07185 0.0230 0.0244 1.0000
11.250 0.7189 0.08414 0.07896 0.0180 0.0246 1.0000
11.500 0.7038 0.09201 0.08685 0.0127 0.0248 1.0000
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Polar data table (+)
Polar graphs
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